The present invention relates generally to gas turbine engines, and, more specifically, to turbine and compressor stator nozzles therein.
A typical gas turbine engine includes a multistage axial compressor through which air is compressed in turn and then mixed with fuel in a combustor and ignited for generating hot combustion gases. The combustion gases flow downstream through corresponding turbines which expand the gases for extracting energy therefrom for powering the compressor, and typically also powering a fan in a turbofan aircraft engine application.
Both the compressor and the turbine include corresponding rows of rotor blades or airfoils extending radially outwardly from supporting rotor disks. Each rotor stage includes a corresponding stator stage defined by an annular nozzle of stator vanes specifically configured for channeling the air for pressurization in the compressor or for channeling the combustion gases for expansion in the turbine.
Although compressors and turbines are functionally different, the corresponding stator nozzles thereof similarly include a row of stator vanes typically mounted from annular outer and inner bands, which in turn are suitably supported to corresponding frames or casings of the engine. Some compressor stators, however, may include solely an outer supporting band, with no inner band being used.
During operation, both the compressor nozzles and turbine nozzles are subject to heating and differential operating temperatures between the outer and inner bands thereof. Air increases in temperature as it is compressed, with the combustion gases having substantially higher temperatures which correspondingly heat the turbine nozzles to even greater temperatures.
Since the stator vanes and supporting bands expand when heated, they are also subject to corresponding thermal growth in diameter, as well as differential radial growth between the outer and inner bands depending upon the particular mode of operation of the engine.
In order to prevent unacceptable restraint in growth of the heated vanes during operation, the supporting outer and inner bands thereof are commonly formed in discrete, arcuate segments for circumferentially interrupting the annular hoop path of the respective nozzles. In this way, the nozzle segments are free to expand and contract relative to adjoining segments for reducing thermally generated reaction stresses during operation.
However, the segmented nozzle bands are subject to leakage through the corresponding splits or gaps therebetween which are commonly sealed for minimizing leakage thereof for maintaining high efficiency of both the compressor and turbine. Typical band seals are in the form of discrete leaf seal elements which are axially disposed in complementary seal grooves formed in the circumferential end faces of the band splitlines.
The resulting construction of the compressor stator nozzles and turbine nozzles include a large number of individual components, including the band seals therefor, and correspondingly increases the complexity of manufacture and cost. However, since the vanes are typically arranged in groups of two or more in each band segment, the segments are readily repairable by removing any one or more of the damaged vane segments and substituting replacement segments.
Accordingly, it is desired to provide an improved nozzle for gas turbine engine compressors and turbines having reduced cost of manufacture and assembly in a simplified construction.
A gas turbine engine nozzle includes outer and inner bands. Each of the bands includes segments circumferentially adjoining at corresponding splits. The splits of the inner band are circumferentially spaced from the splits of the outer band. A plurality of vanes are fixedly joined to the outer and inner segments which collectively define an accordion loadpath therethrough.